XFOIL Version 6.94 Calculated polar for: lb3000 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3134 0.06102 0.04879 -0.0140 1.0001 0.4434 -2.750 -0.2889 0.05907 0.04666 -0.0159 1.0001 0.4510 -2.500 -0.2589 0.05713 0.04432 -0.0197 1.0001 0.4599 -2.250 -0.2349 0.05567 0.04282 -0.0203 1.0001 0.4694 -2.000 -0.2041 0.05425 0.04109 -0.0231 1.0001 0.4809 -1.750 -0.1779 0.05317 0.03997 -0.0239 1.0001 0.4982 -1.500 -0.1512 0.05223 0.03899 -0.0247 1.0001 0.5239 -1.250 -0.1259 0.05134 0.03822 -0.0248 1.0001 0.5564 -1.000 -0.1013 0.05039 0.03753 -0.0244 1.0001 0.6023 -0.750 -0.0826 0.04912 0.03691 -0.0216 1.0001 0.6866 -0.500 -0.0317 0.04631 0.03365 -0.0290 1.0001 0.9999 -0.250 -0.0059 0.04721 0.03361 -0.0301 1.0001 0.9999 0.000 0.0154 0.04814 0.03401 -0.0303 1.0001 0.9999 0.250 0.0354 0.04916 0.03464 -0.0305 1.0001 0.9999 0.500 0.0546 0.05027 0.03546 -0.0308 1.0001 0.9999 0.750 0.0728 0.05151 0.03647 -0.0310 1.0001 0.9999 1.000 0.0901 0.05291 0.03770 -0.0314 1.0001 0.9999 1.250 0.1060 0.05451 0.03917 -0.0319 1.0001 0.9999 1.500 0.1200 0.05638 0.04097 -0.0326 1.0001 0.9999 1.750 0.1313 0.05862 0.04318 -0.0334 1.0001 0.9999 2.000 0.1394 0.06133 0.04589 -0.0346 1.0001 0.9999 2.250 0.1448 0.06446 0.04904 -0.0360 1.0001 0.9999 2.500 0.1497 0.06778 0.05236 -0.0376 1.0001 0.9999 2.750 0.2080 0.07240 0.05685 -0.0494 0.9675 0.9999 3.000 0.2803 0.07677 0.06103 -0.0617 0.9130 0.9999 3.250 0.3191 0.08039 0.06454 -0.0672 0.8819 0.9999 3.500 0.3453 0.08362 0.06768 -0.0703 0.8570 0.9999 3.750 0.3703 0.08683 0.07079 -0.0728 0.8348 0.9999 4.000 0.3964 0.09035 0.07422 -0.0756 0.8196 0.9999 4.250 0.3988 0.09347 0.07730 -0.0755 0.8170 0.9999 4.500 0.4037 0.09663 0.08041 -0.0757 0.8154 0.9999 4.750 0.4099 0.09983 0.08357 -0.0762 0.8153 0.9999 5.000 0.4191 0.10337 0.08705 -0.0773 0.8190 0.9999